A modern industrial gas turbine, as may be used for electrical power generation, may be designed with an annular combustor or an array of can-annular combustors. In the case of a gas turbine with an annular combustor, the combustion chamber is defined circumferentially between the side walls and axially between the inlet plane and the discharge plane. Such a gas turbine is described in commonly assigned U.S. Pat. No. 8,434,313 and is shown in FIGS. 1 through 4. The gas turbine 10, which is shown in detail in FIGS. 1 and 2, has a turbine casing 11 in which a rotor 12 that rotates around a longitudinal axis 27 is housed. A compressor 17, which produces a compressed air flow 2 used for combustion and cooling, is positioned at one end of the rotor 12 and includes blades mounted on the rotor 12. A turbine 13 is arranged downstream of the compressor 17, the turbine 13 also having blades that are mounted on the rotor 12. The compressor 17 compresses air that flows as a compressed air flow 2 into a plenum 14 defined by the turbine casing 11. In the plenum 14, an annular combustor 100 is arranged concentrically around the longitudinal axis 27.
The combustor 100 includes an inner liner shell 33 (proximate to the axis 27) and an outer liner shell 23 (distal to the axis 27), which form the side walls of the combustor 100 and which are radially spaced apart from one another to define an annular interior volume. At the upstream (or head) end of the combustor 100, a front plate 19 spans between the inner liner shell 33 and the outer liner shell 23 to define a combustion zone 15 (sometimes referred to as “zone one”). The front plate 19 defines the inlet plane of the combustion zone 15. Mounted to the front plate 19 at the head end of the combustor 100 is a ring of burners 16, which, for example, may be designed as double-cone burners or EV-burners and which inject a fuel-air mixture into the combustion zone 15. The combustion gases 26 produced by the burners 16 travel from the combustion zone 15 through a transition zone 25 (sometimes referred to as “zone two”) before being discharged from the aft end of the combustor 100 to perform work within the turbine 13. The inner liner shell 33 and the outer liner shell 23 are shaped such that the combustion zone 15 is an annular region of uniform cross-section, while the transition zone 25 defines an annular region of diminishing cross-section to the aft end and discharge plane.
The outer shell 23 and the inner shell 33 are cooled using air 2 from the compressor 17, as discussed below. In order to promote the cooling, an outer cooling shroud 21 is disposed radially outward of the outer shell 23 (that is, distal to the axis 27), thus defining an annular cooling passage 22 between the outer shell 23 and the outer cooling shroud 21. Similarly, an inner cooling shroud 31 is disposed radially outward of the inner shell 33 (that is, toward the axis 27), defining an annular cooling passage 32 between the inner shell 33 and the inner cooling shroud 31. The inner cooling shroud 31 and the outer cooling shroud 21 are connected to the respective inner and outer liner shells 33, 22 by fastening elements 24 (as shown in FIGS. 2 and 4). The inner cooling shroud 31 and the outer cooling shroud 21 may be segmented circumferentially and/or axially (e.g., into upstream cooling shrouds disposed radially outward of the combustion zone 15 and downstream cooling shrouds disposed radially outward of the transition zone 25).
Air 2 from the compressor 17 flows into the cooling passages 22, 32, at the aft end of the combustor 100. Air 2 flows along the liner shells 23, 33 of the combustor 100 in a cooling air flow direction opposite to the direction of the hot gas flow 26 within the combustion zone 15 and the transition zone 25, the air 2 thereby convectively cooling the liner shells 23, 33. At the forward end of the combustor 100, air 2 from the cooling passages 22, 32 is directed into a combustor dome 18 that defines an air plenum 58 from which the air 2 flows into the burners 16 where it mixes with fuel from a fuel line 47. A portion of the air 2 that is directed into the combustor dome 18 flows through the front plate 19, as front plate cooling air 20. The front plate cooling air 20 flows directly into the combustion zone 15.
The inner liner shell 33 and the outer liner shell 23 may be constructed as shell elements or half-shells. When using half-shells, it is desirable for installation and maintenance reasons to secure the half-shells along a parting plane 29 (shown in FIG. 3), which allows an upper half of the shell 23, 33 (e.g., upper half 33a of inner shell 33 in FIG. 3) to be detached from the lower half (e.g., lower half 33b of inner shell 33 in FIG. 3). The parting plane 29 correspondingly has two parting plane welded seams 30, which, in the example of the General Electric GT13E2 gas turbine, are located at the level of the machine axis 27 (i.e., at the 3 o'clock and 9 o'clock positions).
FIG. 4 illustrates a portion of the inner liner halves 33a, 33b, at the parting plane 29 and at the aft end of the annular combustor 100 (that is, forming the tapering portion defining the transition zone 25). The welded seam 30 between the inner liner halves 33a, 33b may be covered with a cooling trough 43 having a plurality of cooling holes (not shown) defined therethrough.
The fastening elements 24, which secure the cooling shroud(s) 31 to the inner liner 33, include a C-shaped bracket 44 and a bolt 45. The bolt 45 is welded or otherwise affixed (optionally, with a washer) to the center portion of the C-shaped bracket, and the respective ends of the bracket 44 are welded or otherwise affixed to the outer surface of the inner liner half 33a, 33b. The fastening elements 24 are aligned along a common plane or axis 49 from the forward end of the inner liner half 33a, 33b to the aft end of the inner liner half 33a, 33b. The cooling shrouds 31 are disposed over the fastening elements 24 and are secured thereto by a threaded nut 46 (shown in FIG. 2), optionally with a washer.
The inner and outer liner shells 33, 23 of the gas turbine 10 are known to be thermally and mechanically highly stressed during operation. The strength properties of the material of the shells 23, 33 are greatly dependent upon temperature. In order to keep the material temperature below the maximum permissible material temperature level, the shells 23, 33 are convectively cooled, as described above. One challenge to be overcome in the design of the cooling shrouds 21, 31 is the accommodation of thermal expansion, which occurs during the operation of the gas turbine 10. Another challenge to be overcome in the design of the cooling shrouds 21, 31 is the reduction of vibrations of the cooling shrouds 21, 31, as may be expected to occur during the operation of the gas turbine 10, which may negatively impact the part life and shorten the maintenance intervals of the combustor 100.